A gas turbine engine typically comprises at least one combustor, which mixes air from a compressor with a fuel. This fuel and air mixture combusts after being introduced to an ignition source. The resulting hot combustion gases pass through the combustion system and into a turbine, where the gases turn the turbine and associated shaft. A gas turbine engine is most commonly used for either propulsion for propelling a vehicle or harnessing the rotational energy from the engine shaft to drive a generator for producing electricity. Most land-based gas turbine engines employ a plurality of combustors arranged in a can-annular layout around the engine. Referring to FIG. 1, a representative land based gas turbine engine 10 of the prior art is shown in partial cross section. Gas turbine engine 10 comprises an inlet region 11, an axial compressor 12, a plurality of combustors 13, each in fluid communication with a transition duct 14, which are in fluid communication with a turbine 15. The hot combustion gases drive the turbine, which turns shaft 17 before exiting through outlet 16. Shaft 17 is coupled to the compressor, and for power generation, to an electrical generator (not shown).
The operating temperatures of the combustors 13 are typically well over 3000 degrees Fahrenheit, while the temperature limits of the materials comprising combustors 13 are much lower. Therefore, in order to maintain the structural integrity for continued exposure to the hot combustion gases, combustors 13 are cooled, typically by air from compressor 12. However, it is critical to only use the minimal amount of cooling air necessary to lower the operating metal temperatures of combustor 13 to within the acceptable range, and not use more air than necessary nor allow any cooling air leakage.
In order to maximize the efficiency of the gas turbine engine, it is imperative to minimize any leakage of air from compressor 12 that is not intended for cooling combustors 13, such that all air not intended for cooling, passes through combustors 13 and undergoes combustion. Leakage areas are especially common between mating components such as the interface region between combustor 13 and transition duct 14. Seals or tight tolerances between such mating components are typically employed to overcome such leakages that can reduce overall performance and efficiency. However, it is also imperative to provide adequate cooling to an interface region.
Examples of prior art seals and cooling designs for the interface region between combustor 13 and transition duct 14 are disclosed in U.S. Pat. Nos. 5,724,816 and 6,334,310. The '816 patent pertains to a plurality of axial channels that are formed between an inner member and an outer member and can be used to cool the aft end section of a combustion liner where it interfaces with a transition duct. An example of this configuration is shown in FIG. 2 where a combustion liner is provided having a plurality of axial cooling channels 18. The '310 patent pertains to an alternate manner to cool this same region of a combustion liner and can be used in conjunction with the prior art combustion liner shown in FIG. 2. Specifically, a combustion liner includes an outer cooling sleeve that contains a plurality of cooling holes 19 for supplying cooling air to the region between the liner and the outer cooling sleeve. The outer cooling sleeve includes a swaged end such that when the outer cooling sleeve is welded to the combustion liner the stresses imparted to the outer cooling sleeve by a transition duct are moved away from the weld joint. Often times these combustion liners are also accompanied by at least one spring seal for sealing against the inner wall of a transition duct.
While each of these designs are directed towards providing adequate cooling at the interface region of a combustion liner and transition duct, improvements can be made such that cooling effectiveness is improved, extending component life, while simultaneously minimizing unnecessary cooling air leakage.